Wing tip device attachment apparatus and method

ABSTRACT

An aircraft wing subassembly including: a wing skin defining a first outer surface, and, a structural reinforcement member, the structural reinforcement member defining a second outer surface, wherein the structural reinforcement member is arranged within the wing such that the first outer surface and the second outer surface form part of an outer wing surface.

RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No. 15/295,262 filed Oct. 17, 2016, which is a continuation of U.S. patent application Ser. No. 13/317,784 (now U.S. Pat. No. 9,499,255) filed Oct. 28, 2011, and claims priority to GB Application No. 1018185.7, filed 28 Oct. 2010, the entire contents of these applications are incorporated by reference.

BACKGROUND

The present invention is concerned with wing tip device attachment apparatus and method. More particularly, the present invention is concerned with an apparatus and method for attaching a wing tip device such as a winglet to the tip of a passenger aircraft wing.

Wing tip devices are well known in the art. Devices such as winglets, raked wing tips and fences are collectively known as aerodynamic wing tip devices and are used to reduce the effects of lift induced drag.

Lift induced drag is caused by the generation of vortices at the wing tip. Such drag is mitigated by an increase in wing span. Increases in wing span in the plane of the wing are not always possible due to space requirements at, e.g. airports. As such out-of-plane extensions to the wing are commonly used to increase the effective wing span without increasing the geometric span of the aircraft. These take the form of aerodynamic wing tip devices.

There is an ever increasing drive to increase the efficiency of passenger aircraft. One way to achieve increased efficiency is to increase the size of the aerodynamic wing tip device. Typical ratios of winglet span to the thickness of the wing tip at the attachment point (wingbox thickness) are commonly above 10 and may be as high as 12 to 15 in modern passenger aircraft. Because the thickness of the wing is low at the tip, the vertical moment arm available to react to the loads generated by the wing tip device under both its own weight and under aerodynamic forces is low. Therefore, the forces and stresses generated in this area are high.

Other wing tip devices include external tanks, refuelling pods etc whose primary purpose is not to improve the aerodynamic efficiency of the wing, but nerveless are attached to, and produce a force on, the wing tip.

Known wing tip devices are generally attached in one of two manners. The first is to use a series of splice plates or butt straps which span the upper and lower skins of the wing tip device and the wingbox at the point at which they join.

The second method is to use abutting plates joined by tension bolts.

Disadvantageously, both of these methods only utilise a very small moment arm to react the loads. The splice plates transfer load through the wing skin, which is primarily designed to absorb the bending load across the wing span, and less well suited to absorb local, concentrated, loads. As such, the large local load introduced into the wing skin requires structural reinforcement. The additional weight this causes is undesirable.

Further, because the wing skin is not particularly strong, many joining locations are required to spread the applied load. Although this has the desired effect of reducing the load per joining location, it creates a statically indeterminate system making the loads at each point difficult to predict. Therefore each joint is typically over-engineered adding weight and cost to the aircraft.

Cyclic loading is common in aircraft. This introduces additional structural requirements, in particular to the tension bolt design. In order to mitigate the effects of fatigue, the bolts have to be pre-tensioned with an interference fit. This is undesirable as it adds complexity to the manufacturing process, and makes maintenance and replacement more difficult

A still further problem with the above two methods is that because of the heavy bolting and large surface area of contact between the various components in both methods, the interface between the parts is quite sensitive to differences in geometry at the interface. As such, any mis-match between the two components needs to be addressed with fettling upon assembly. This increases the cost of assembly and makes it more difficult to replace the wing tip devices in service.

Further, temperature effects and loading in use can cause differential expansion/contraction of the wing tip device and the wing tip, which can cause high stresses at the mounting points.

SUMMARY OF INVENTION

It is an aim of the present invention to overcome or at least mitigate one or more of the above problems.

According to a first aspect of the invention there is provided a wing tip device for attachment to a wing tip of a powered aircraft, the wing tip device comprising: a first mounting formation, a second mounting formation spaced apart in a spanwise direction relative to the first mounting formation, a third mounting formation spaced apart in a chordwise direction relative to the first and second mounting formations, wherein each of the first, second and third mounting formations are configured for attachment to at least one of a wing spar and a wing rib, and, at least one of the mounting formations is configured so that relative movement in the spanwise direction of the wing tip device to a wing tip is permitted in use.

Advantageously, thermally induced spanwise relative expansions and contractions between the wing tip device and wing tip do not cause excessive stresses because of the relative movement permitted therebetween. Similarly, relative movement resulting from flexing of the wings is permitted without causing significant stresses.

Further, by designing the system such that certain attachment points are not required to react the incident forces, they can be designed as such, increasing the predictability of force magnitude on the other attachment points, and moving the system towards static determinancy.

Preferably the mounting formations are arranged such that relative movement is permitted at least two of the mounting formations to provide a statically determinate loading system. More preferably another of the mounting formations is configured so that relative movement in the chordwise direction of the wing tip device to the wing tip is permitted, and still another of the mounting formations is configured so that relative movement in the vertical direction of the wing tip device to the wing tip is permitted.

Wing spars are load-bearing components and if constructed from composite, also tend to be over-engineered (i.e. thicker than necessary) in the region of the wing tip because they are not easily tapered from the fuselage.

The present invention utilises this material to react the loads from the winglet. Further, the rear spar extends across the entire span of the wing, and as such a large moment arm can be incorporated into the design to reduce the local loads and stresses.

Still further, given the mounting formation arrangement prescribes a statically determinate assembly, it will be known what loads (type and magnitude) will be felt where, thus meaning that the structures can be more efficiently engineered, without the need to account for uncertainties in the amount of load reacted by a given component.

According to a second aspect of the invention there is provided a wing tip device for a powered aircraft comprising a mounting member extending from the wing tip, the first mounting member defining a first attachment formation comprising a first pivotable joint means for rotation about a first axis at a first position proximate the wing tip device, and defining a second attachment formation distal to the wing tip device.

Such a structure permits rotation of the wing tip device into position about the first pivotable joint means and subsequent attachment at the second attachment point. It will be noted that by “pivotable joint means” we mean, inter alia, a pivot shaft, pivot bore or pivot assembly.

According to a third aspect of the invention there is provided a method of assembling a wing tip device to a wing of a powered aircraft comprising the steps of providing a wing tip, providing a wing tip device, pivotably attaching the wing tip device to the wing at a first position, pivoting the wing tip device about the first position, attaching the wing tip device to the wing at a second position spaced from the first position.

Advantageously, the method permits easy installation of wing tip devices upon assembly and in service.

According to a fourth aspect of the invention there is provided an aircraft wing subassembly comprising a wing skin defining a first outer surface, and a structural reinforcement member within the wing, the structural reinforcement member defining a second outer surface, which structural reinforcement member is arranged within the wing such that the first outer surface and the second outer surface form part of an outer wing surface.

Advantageously, by using the reinforcement as part of the wing skin, the reinforcement can be made as large as possible, thus providing maximum resistance to bending moments.

SUMMARY OF DRAWINGS

A wing tip device attachment apparatus and method in accordance with the invention will now be described by way of example and with reference to the accompanying figures in which:

FIG. 1 is a front view of a wing tip and wing tip device;

FIG. 2 is a free body diagram of a part of the wing tip and wing tip device of FIG. 1;

FIG. 3a is a view of a wing tip device and mounting apparatus in accordance with the present invention;

FIG. 3b is a schematic view of the wing tip device and mounting apparatus of FIG. 3 a;

FIG. 4a is a perspective view of a first constraint type;

FIG. 4b is a perspective view of a second constraint type;

FIG. 4c is a perspective view of a third constraint type;

FIG. 4d is a perspective view of a fourth constraint type;

FIGS. 5a to 5f are schematic plan views of various reaction loads exhibited by the mounting apparatus of FIG. 3;

FIGS. 6a to 6d are front schematic views of an attachment method of a wing tip device according to the present invention;

FIG. 7 is a perspective view of a second wing tip device in accordance with the present invention;

FIG. 8a is a chordwise cross-section view of a wing tip fitted with the wing tip device of FIGS. 7; and

FIG. 8b is a close-up view of a part of the wing tip device of FIG. 7 shown in plan section view.

DETAILED DESCRIPTION OF INVENTION

Referring to FIG. 1, a wing assembly 100 comprises a wing tip 102 (formed by the end of a wing) and a wing tip device 104. In this example the wing tip device 104 is a winglet. The wing tip 102 terminates at an outboard end 106 with thickness Tw. This area is also known as the wingbox. The wing tip device 104 extends from the outboard end 106 and has a winglet span Sw1. Turning to FIG. 2, the wing tip device 104 experiences an aerodynamic lift force Fw1 which acts through its centre of pressure 108 generally towards the aircraft fuselage (not shown).

The wing tip device 104 is attached to the wing tip 102 at the outboard end 106. As such a torque, Tw1, is generated which is a product of the winglet force Fw1 and the perpendicular distance Lw1 to the centre of the outboard end 106 of the wing tip 102 (also known as the winglet moment arm).

In order to keep the wing tip device 104 stably attached to the outboard end 106 of the wing tip 102, the torque Tw1 created by the winglet force Fw1 must be reacted at the outboard end 106. Because the moment arm available at the outboard end 106 can only be as high as the wingbox thickness Tw, the reaction forces Fw1, Fw2 are extremely high. As such the material in the area of the outboard end 106 of the wing tip 102 has to be reinforced adding weight and complexity to the aircraft.

As mentioned above, known attachment methods include splice plates which span the upper and lower skin of the wing tip device 104 and the wing tip 102. Alternatively abutting perpendicular plates at the outboard end 106 which are used and held in position by tension bolts. In both cases a moment arm defined vertically between the two wing covers is used to react the forces.

Turning to FIGS. 3a and 3b , a wing assembly 110 is shown comprising a wing tip 112 and a winglet 114. The wing tip 112 comprises a front spar 116 and a rear spar 118, both running in a spanwise direction and converging along the span of the wing towards the wing tip. A series of ribs 120 are positioned along the wing span and extend in the direction of the wing chord. A wing tip end 122 is provided at the end of the wing tip. The spars and ribs are covered by a wing skin 124 primarily designed to present an aerodynamic surface to the airflow. The wing tip 112 terminates at an outboard end 126.

The winglet 114 comprises a winglet root 128 and a free end 130 distanced from and vertically spaced from the winglet root 128.

A main beam 132 extends from a position partway between the free end 130 and the winglet root 128 and extends towards the winglet root 128 and beyond into the wing tip 112 as will be described below. The main beam 132 is spaced towards the rear of the winglet 114. A canted spar 134 runs from the position midway along the winglet 114 towards the winglet root 128 but diverges from the main beam 132 towards the forward part of the winglet 114. The canted spar 134 extends into the wing tip 112 as will be described below.

The main beam 132 and the canted spar 134 are supported by a number of winglet ribs 136 which extend chordwise within the winglet 114. A winglet skin 138 covers the winglet in order to present an aerodynamic surface to the airflow.

Referring to FIG. 3b , the wing assembly 110 is shown in schematic form. The main beam 132 extends from the winglet 114 into the wingbox of the wing tip 112 to be generally parallel and adjacent to a forward face of the rear spar 118. The main beam 132 is attached to the rear spar 118 at two positions; position B proximate the first rib 120 and position A proximate the wing end rib 122. It will be noted that A and B are spaced apart and, in particular, spaced apart by a distance which is larger than the thickness of the wingbox Tw.

The canted spar 134 also extends into the wing tip 112, but in this example is only arranged to abut the wing tip rib 122 and is attached thereto at point C.

A, B and C are therefore first, second and third mounting formations, and will be described in greater detail below.

Turning to FIGS. 4a to 4d , various examples of attachment methods are shown. FIG. 4a shows the main beam 132 attached to the outermost rib 120 via a spigot joint 140.

FIG. 4b shows the main beam 132 attached to the rear spar 118 via a single lap shear joint 142.

FIG. 4c shows a main beam 132 passing through the wing tip rib 122 and attached to the rear spar 118 and a sub-spar 144 extending from the wing tip rib 122 via a double lap shear joint 146.

FIG. 4d is a plan view of a canted spar 134 attached to the front spar 116 via a web in single lap shear

In the examples shown in FIG. 3b , the attachment at B is a spigot joint per FIG. 4a , the attachment at A is a double lap shear joint as shown in FIG. 4c and the attachment at C is a locking pin.

Referring to FIGS. 5a to 5f , reaction of the various loads and moments will be described. In all cases the co-ordinate system used is X in a positive spanwise direction, Y in a fore aft (chordwise) direction and Z in the vertical direction.

Referring to FIG. 5a , the bending moment MY (the tendency of the tip of the wing tip device to move towards the fuselage) is reacted at attachment points A and B as shown. This is the type of bending moment produced by the winglet lift force Fw1, and as shown it is reacted over a large moment arm (the distance from A to B), reducing the force, and hence stress levels.

Referring to FIG. 5b , the moment MZ (mainly resulting from drag) is reacted where the winglet abuts the wing tip at the trailing edge proximate the attachment point A and at the connection at point C which is held in direction X.

Referring to FIG. 5c , the torsion MX is reacted by the fact that attachment points A and C are constrained in the vertical direction (i.e. in direction Z).

Referring to FIG. 5d , FY (drag force) is reacted primarily at attachment point C.

Referring to FIG. 5e , FZ (lift) is reacted at points A and C.

Finally, referring to FIG. 5f , the side force FX is reacted at points A and C.

The release of certain degrees of freedom (e.g. the inability of the spigot at B to react the side force FX) allows the system some relative movement to avoid thermally induced stresses whilst making the loads more predictable (moving towards a statically determinate system). For example, because the joint at point B does not need to react the side force, it can be made smaller as a result (i.e. can be optimised for a more predictable load case).

It will be noted that because the present invention only uses three attachment points, it is possible to constrain the winglet 114 in a manner which makes the system statically determinate. Therefore, each attachment point can be designed around a known load case. This offers an advantage over the prior art in which generally a high number of fixings are used for load-bearing purposes and consequently a statically indeterminate system is formed in which the exact load case on each attachment point is unknown. Therefore each attachment point has to be over-engineered to cope with the worst possible case.

Referring to FIGS. 6a to 6d , a method of attachment of a wing tip device is shown. A wing assembly 200 is shown comprising a wing tip 202 and a winglet 204. The winglet 204 is attached to the wing tip 202 in a similar manner to the wing assembly 110. In the wing assembly 200, the attachment points A and C have their horizontal chordwise axes (parallel to axis Y) of rotation aligned as will be described below. The winglet 204 is moved proximate the wing tip 202 on a trolley jack or similar as shown in FIG. 6a . The jack 206 is elevated to move the winglet 204 such that the attachment points A, C are aligned with their respective receiving formations on the wing tip 202. The joints can then be assembled such that the winglet 204 can be rotated above the axis Y at the attachment points A, C from the position shown in FIG. 6B to the position shown in FIG. 6C. The winglet 204 is rotated into place and the attachment point B is secured in order to prevent any rotation of the winglet 204 relative to the wing tip 202. The jack 206 can then be removed, as shown in FIG. 6d .

This method of assembly demands an interruption in the skin on the top of the wing tip 202. This can be achieved by making the winglet mean beam part of the aerodynamic surface of the wing (see below) or providing a replaceable panel in the wing skin. The method permits replacement of the winglet in-field without the need for an overhead crane and/or hanger space.

Referring now to FIG. 7, a wing assembly 300 is shown comprising a wing tip 302 and a winglet 304.

The wing tip 302 comprises a front spar 306 and a rear spar 308. A front spar 306 comprises two flanges extending in a chordwise direction; an upper flange 310 and a lower flange (not visible). The flanges extend towards the rear spar 308. Similarly, the rear spar 308 comprises an upper flange 312 and a lower flange 314 both of which extend towards the front spar 306. A rib 316 extends between the spars 306, 308 in a chordwise direction at the widest parts of the flanges 310, 312, 314.

The winglet 304 comprises a flat main beam 318 which extends substantially parallel to the skin of the winglet 304. The main beam 318 tapers from a point midway along the winglet 304 to its thickest cross-section at a mid-point 320 at the position where the winglet and the wing tip meet and tapers inwardly again at attachment point 322 within the wing tip 302.

The beam 318 is attached to the wing tip 302 via a spigot at point B, a lap shear joint at point A and a further lap shear joint at point C. The axes of rotation of the lap joints at A and C are aligned such that the winglet 304 can be assembled to the wing tip 302 in a similar manner as described in FIGS. 6a to 6 d.

It will be noted that the beam 318 tapers from the point of maximum bending moment at area 320 to areas of lower bending moment at its opposite ends within both the winglet 304 and the wing tip 312. Referring to FIG. 8a , a section is shown through the wingbox of the wing tip 302 proximate points A and C. It will be noted that the main beam 308 is designed to form part of the aerodynamic surface of the wing. Referring to FIG. 8b , the skin of the wing tip 302 is transitioned to the main beam 318 by use of flexible skin panels 324 which define a tapered region 326 to allow transition to the surface of the beam 318. In this way, the beam 318 can be as large as possible in order to provide as much area for reaction on the various loads and stresses it undergoes.

It will also be noted that by making the beam 318 part of the wing skin, the assembly process as shown in FIGS. 6a to 6d is made easier as the skin of the wing tip 302 does not need to be replaced over the beam 318.

Variations fall within the scope of the present invention. 

The invention is:
 1. An aircraft wing subassembly comprising: a wing skin defining a first outer surface, and, a structural reinforcement member, the structural reinforcement member defining a second outer surface, wherein the structural reinforcement member is arranged within the wing such that the first outer surface and the second outer surface form part of an outer wing surface.
 2. The aircraft wing subassembly according to claim 1, in which the structural reinforcement defines a third outer surface on an opposite side of the wing to the first outer surface and the structural reinforcement member is arranged within the wing such that the first outer surface and the third outer surface form part of an outer wing surface.
 3. The aircraft wing subassembly according to claim 1, in which the wing skin tapers towards the second outer surface to form a smooth transition therewith.
 4. The aircraft wing subassembly according to claim 3, in which the taper is defined on a flexible member bridging the wing skin and the second surface.
 5. The aircraft wing subassembly according to claim 2 in which the structural reinforcement defines a rectangular cross section with the second and third outer surfaces being defined on opposite sides thereof.
 6. An aerodynamic structure for an aircraft comprising: a first wing skin defining a first portion of a first outer aerodynamic surface of the aerodynamic structure; a main beam spanning between the first outer aerodynamic surface and a second outer aerodynamic surface of the aerodynamic structure; and an outer surface of the main beam forming a second portion of the first outer aerodynamic surface.
 7. The aerodynamic structure of claim 6 wherein the outer surface is a first outer surface and the main beam includes a second outer surface opposite to the first outer surface and the second outer surface forms a portion of the second outer aerodynamic surface of the aerodynamic structure.
 8. The aerodynamic structure of claim 7 wherein the main beam includes sides spanning the first and second outer surfaces.
 9. The aerodynamic structure of claim 8 wherein a front side of the sides of the main beam is attached to a leading edge structure of the wing and a rear side of the sides is attached to a trailing edge structure of the wing.
 10. The aerodynamic structure of claim 6 wherein the main beam is hollow.
 11. The aerodynamic structure of claim 6, wherein a thickness of the first wing skin tapers towards and overlaps the main beam to form a smooth transition surface between the first wing skin and the main beam.
 12. The aerodynamic structure of claim 11, wherein the taper in the first wing skin is a flexible member.
 13. The aerodynamic structure of claim 6, wherein the aerodynamic structure is a wing.
 14. The aerodynamic structure of claim 13, wherein the wing includes a winglet, and the main beam spans between the winglet and a fixed wing of the wing.
 15. The aerodynamic structure of claim 6 wherein the outer surface of the main beam is adjacent an edge of the wing skin. 